The present invention relates to gas turbine engines having a multi-stage turbine. More particularly, the present invention relates to an interstage sealing and torque ring for disposition between adjacent stages in a gas turbine engine having a multi-stage turbine.
Gas turbine engines having multiple turbine stages include sealing arrangements between adjacent stages for improved operating efficiency. The sealing arrangements are directed to confining the flow of hot combustion gases to flow in an annular path around and between the stationary turbine stator blades, or nozzles, and also around and between the adjacent rotor blades. In addition to serving to confine the hot combustion gases to flow in a predetermined annular path, such sealing arrangements also serve to confine and to direct cooling air that is provided to cool the turbine disks, the turbine blade roots, and also the interior of the rotor blades themselves. In that regard, providing rotor blade cooling passages allows higher turbine inlet temperatures, which results in higher thermal efficiency of the engine and higher power or thrust output. However, the air for cooling the turbine rotor blades must be suitably confined and directed so that it is not dissipated by passing into the hot gas stream, but, instead, passes over and through the surfaces and structures intended to be cooled, and so that it also passes into the rotor blade internal passageways to provide the desired rotor blade cooling effect.
In the past, various turbine interstage seal configurations have been proposed. Some of the previous arrangements incorporated rotating disks that define sealing rings that include a peripheral labyrinth seal for engagement with a stationary, annular sealing ring carried on an interior surface of the stationary turbine nozzle. Some arrangements included forwardly and rearwardly-extending arms for engagement with rotor blade retaining rings carried by the adjacent turbine rotor stages. Other previous arrangements included integrally-formed rotor blade retainers to prevent axial movement of the rotor blades relative to the rotor disks, and inner hubs that connected with one or more of the adjacent rotor disks for causing the seals to rotate with the turbine rotor. However, multiple element seal and blade retainer arrangements involve additional parts and additional assembly operations, and they can also involve problems of maintaining proper orientation of the several parts to provide the desired sealing effect.
In addition to the logistical and assembly problems that inhere in such multiple-element arrangements, it is also necessary that the sealing arrangement be capable of accommodating axial and radial movements of the turbine stage elements during engine operation. The several elements are subjected to a range of different loadings and different rates of expansion based upon local part temperatures and also based upon engine and aircraft operating conditions.
Accordingly, it is desirable to provide a turbine interstage sealing member that can provide an effective seal to confine combustion gases to flow in a desired annular channel, to separate combustion gases from cooling air flows, and to do so with a unitary structural arrangement that facilitates assembly of a multi-stage turbine and that is also capable of responding to thermally- and mechanically-induced size and orientation changes of the turbine structural elements.
Briefly stated, in accordance with one aspect of the present invention, an interstage sealing ring is provided for sealing a space between adjacent turbine rotors of a multi-stage turbine. The sealing ring includes a substantially disk-shaped body member having first and second axially-spaced, substantially radially-outwardly-extending arms. The first radial arm includes a forwardly-extending, substantially axial first arm that terminates in an engagement surface for engagement with the surface of a first turbine rotor. The second radial arm. includes a rearwardly-extending, substantially axial second arm that terminates in an engagement surface for engagement with a second turbine rotor that is spaced axially from the first turbine rotor. The sealing ring defines a bridging member between adjacent turbine rotors to confine cooling air that is provided in the space between the adjacent turbine rotors so the cooling air flows into cooling air passageways within the turbine rotor blades.
In accordance with another aspect of the present invention, a turbine cooling system is provided for a multi-stage turbine that includes a pair of axially-spaced turbine rotors. The cooling system includes an interstage cooling chamber defined by a space between a pair of axially-adjacent turbine rotors, and an annular rotor-disk-connecting ring positioned adjacent a hub of each turbine rotor and interconnecting the adjacent rotor disks for co-rotation. An interstage sealing member is positioned radially outwardly of the disk connecting ring. A plurality of cooling air passageways are provided in an upstream turbine disk for conveying cooling air through the disk and into the interstage cooling chamber. The interstage sealing member includes a substantially disk-shaped body member having a pair of axially-spaced, substantially radially-outwardly extending arms. A forwardly-extending, substantially axial first arm terminates in an engagement surface for engagement with a first turbine rotor, and a rearwardly extending, substantially axial second arm terminates in a engagement surface for engagement with a second turbine rotor that is spaced axially from the first turbine rotor.